The present invention generally relates to a gas turbine engine, and more particularly to a cooling scheme for a gas turbine engine component.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. Air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to add energy to expand the air and accelerate the airflow into the turbine section. The hot combustion gases that exit the combustor section flow downstream through the turbine section, which extracts kinetic energy from the expanding gases and converts the energy into shaft horsepower to drive the compressor section.
The turbine section of the gas turbine engine typically includes alternating rows of turbine vanes and turbine blades. The turbine vanes and blades typically include at least one platform and an airfoil which extends from the platform. The turbine vanes are stationary and function to direct the hot combustion gases that exit the combustor. The rotating turbine blades, which are mounted on a rotating disk, extract the power required to drive the compressor section. Due to the extreme heat of the hot combustion gases that exit the combustor section, the turbine vanes and blades are exposed to relatively high temperatures. Cooling schemes are known which are employed to cool the platforms and the airfoils of the turbine vanes and blades.
For example, impingement platform cooling and film cooling are two common methods for cooling the platforms and airfoils of the turbine vanes and blades. Both methods require a dedicated amount of air to cool the platform. Disadvantageously, there is often not enough cooling airflow available to supply both the airfoil and the platforms with a dedicated airflow.
In addition, both impingement platform cooling and film cooling require holes to be drilled through the platforms to facilitate the dedicated airflow needed to cool the platform. The holes may be subject to hot gas ingestion due to insufficient backflow margin. Insufficient backflow margin occurs where the supply pressure of the cooling airflow is less than that of the hot combustion gas path. Where this occurs, hot gas ingestion may result (i.e., hot air from the hot combustion gas path enters the cooling passages of the turbine vanes and blades through the cooling holes) thereby negatively effecting the cooling benefits provided by the cooling holes. Further, even if the cooling air supply pressure is sufficient, the drilled cooling holes may cause undesired aerodynamic losses.
Accordingly, it is desirable to provide an improved cooling scheme for a gas turbine engine component which provides efficient and simultaneous cooling of an airfoil and a platform of the gas turbine engine component.